This invention relates generally to gas turbine engines and in particular to flowpath structures within a gas turbine engine.
A typical gas turbine engine includes a turbomachinery core having a high pressure compressor, a combustor, and a high pressure turbine in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. In practical applications the core is typically combined with other elements such as power turbines, fans, augmentors, etc. to create a useful engine for a specific application, such as turning a propeller, powering an aircraft in flight, or driving a mechanical load.
Gas turbine engines include a flowpath defined in part by ducts, liners, tubes, and similar structures that directs a working fluid through the various components of the engine. Some portions of the flowpath are subject to hot, high-velocity gases. Prior art flowpath components, particularly those in the hot section of the engine, often use metal alloy structures protected with a thermal barrier coating (“TBC”).
Metallic structures can be replaced with materials having lower density, such as ceramic matrix composites (CMCs). Such materials offer significant weight savings compared to metal alloys.
One problem with CMC materials is that they cannot be fabricated or mechanically fastened in the same way as components made from metal alloys, and therefore cannot be substituted directly for metallic components.
Another problem with CMC materials is that they have relatively low tensile ductility or low strain to failure when compared to metallic materials. Also CMCs have a coefficient of thermal expansion (CTE) significantly different from metal alloys.